Gas turbine engines are generally known in the art and used in a wide range of applications, such as aircraft engines and auxiliary power units for aircraft. In a typical configuration, the turbine of such engines includes rows of stator vanes and rotor blades disposed in an alternating sequence along the axial length of a generally annular hot gas flow path. The rotor blades are mounted at the periphery of one or more rotor disks that are coupled in turn to a main engine shaft. Hot combustion gases are delivered from an engine combustor to the annular hot gas flow path, thus resulting in rotary driving of the rotor disks to provide an engine output.
In most gas turbine engine applications, it is desirable to regulate the operating temperature of certain engine components in order to prevent overheating and potential mechanical failures attributable thereto. That is, while the engine stator vanes and rotor blades are specially designed to function in the high temperature environment of the mainstream hot gas flow path, other engine components may not be designed to withstand the high temperatures of the mainstream hot gas flow. Accordingly, in many gas turbine engines, the volumetric space disposed radially inwardly or internally from the hot gas flow path includes an internal engine cavity through which a cooling air flow is provided. The cooling air flow is normally obtained as a bleed flow from a compressor or compressor stage forming a portion of the gas turbine engine. The cooling of the internal engine cavity attempts to maintain the temperatures of the rotor disks and other internal engine components that are suitable for their material and stress level.
In many conventional engines, relatively high cooling air flows have been used to obtain satisfactory temperature control of engine components within the cooled internal engine cavity. In addition, the demand for cooling flow has been impacted by an irregular and unpredictable ingestion of mainstream hot gases from the hot gas flow path into the internal engine cavity. Various attempts to prevent flow between adjacent stator vanes and rotor blades have primarily involved the use of overlapping lip-type structures in close running clearance, often referred to as flow discouragers, but these structures have not been as effective as desired in preventing hot gas ingestion.
A variety of baffle-type structures and techniques have also been proposed, in addition to the traditional flow discouragers, in effort to minimize hot gas ingestion into the internally cooled cavity of gas turbine engines. Such approaches have included pockets with complex shape, some of which receive separate flows of cooling gas, to prevent hot gas ingestion. In the past, these techniques may have been less effective than desired, and/or may have used structures of complex shape and/or mounting arrangements at the time of initial engine production.
Accordingly, it is desirable to provide an improved gas turbine engine assembly that reduces or eliminates the effects of hot gas ingestion. In addition, it is desirable to provide a recirculation cavity that captures and recirculates ingested hot gas with high efficiency. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.